This invention relates to correcting undesirable flight characteristics of an aircraft. More particularly, this invention relates to an aerodynamic device used to remedy a nonlinear, unstable high Mach number stall characteristic of a swept-wing jet transport commonly called "pitch up".
As known to those skilled in the art, air flowing over the upper surface of an airplane wing separates from the wing surface when the wing reaches a given angle of attack at sufficiently high airspeed, resulting in a loss of lift or a condition known as stall. Moreover, with a swept-wing airplane operating at high Mach numbers, e.g., above approximately 0.6, as the angle of attack of the wing is increased, air flowing over the upper surface of the outboard wing separates from the wing before air flowing over the upper surface of the inboard separates; thus, the outboard section of the wing effectively stalls at a lower angle of attack than the inboard section.
Two factors contribute to the earlier separation of airflow over the outboard wing. The first factor is spanwise airflow over the wing due to the sweep of the wing. The spanwise airflow augments the adverse pressure gradient experienced by air flowing over the outboard wing's upper surface, thereby contributing to earlier separation of airflow from the outboard wing's upper surface. A second factor causing earlier separation of airflow over the outboard wing is shock-induced separation. As air flowing over the outboard wing's upper surface reaches supersonic velocity, a shock is formed, and it is formed at lower airspeeds than on the inboard wing. The shock takes energy out of the airflow behind it, causing the airflow to tend to seperate from the wing's surface. Thus, as the angle of attack of a swept wing is increased when the airplane is operating at high Mach numbers, the outboard wing loses lift before the inboard wing.
On present-day swept-wing jet transports, the pitch axis of the airplane extends laterally through the center of gravity of the airplane and is generally located at approximately 25% of the mean aerodynamic chord of the wing. The centers of lift of the inboard and outboard wings are generally located rearward of the pitch axis; thus, the lifting forces generated by the inboard and outboard wings create a pitching moment that tends to force the nose of the airplane downwardly. In a stable flight condition, the pitching moment created by the lift of the wings is counteracted by a downward force exerted by the airplane's horizontal stabilizer. When the angle of attack of the wing is increased to the point where separation occurs on the outboard wing causing a decrease in the lift component, a resultant tendency toward pitch up occurs since the lift component of the inboard wing is still linearly increasing with increasing angle of attack and the downward force of the horizontal stabilizer increases proportionally with increasing inboard wing lift. The unstable character of this phenomenon is undesirable and is further aggrevated by unpredictable factors such as local atmospheric conditions including local ambient wind patterns, e.g., gust upsets. Because of the abruptness of the effect at certain Mach numbers and the element of unpredictability, combined with the associated increase in load factor and the increase in buffet load at high Mach numbers and high angles of attack, various aerodynamic solutions to the problem have been proposed that are not dependent upon pilot response.
Two basic approaches have been used attempting to solve the high speed pitch up problem. In one approach, devices have been employed to enhance the outboard wing lifting capability. These include fences on the outboard wing or midwing, wing twist, profile camber changes, leading edge contour changes, wing planform changes (saw-tooth, gloves, etc.) and vortex generators located at the midspan of a wing. However, none of these have completely eliminated the pitch up problem. Most have increased the coefficient of outboard wing lift as a function of angle of attack but have not eliminated the pitch up loop nor changed its magnitude.
For example, midspan vortex generators provide a solution using this approach in that they were found to correct flow problems in the midspan area and prevent spanwise flow from affecting the outboard wing up to a certain angle of attack, thus helping airplane stability and extending the linear part of the pitching moment curve. However, this contribution to extension of the stable part of the pitching moment curve occurs at Mach numbers greater than the cruise Mach number of present day jet transports. In all cases, the solution with a vortex generator is likely to be very dependent upon the configuration of the wing, or local flow condition and shock position on the outboard wing. Thus, identical placement of voxtex generators on wings having different configurations are unlikely to achieve the same results.
The second basic approach used in attempting to solve the high speed pitch up problem is the reduction of inboard wing lift coincidentally with the loss of lift on the outboard wing. Efforts concentrating on this approach are relatively recent. One prior art device using this approach is an extendable/retractable discontinuity on the upper surface of the inboard wing and forward of the 25% chord. This discontinuity can be a stall strip or a leading edge spoiler progressivey extended with increasing Mach number and angle of attack. This type of device has been found to be effective in correcting high speed pitch up, however, it has several drawbacks. One is a sudden leading edge separation at high Mach numbers and angles of attack resulting in a very rapid rise in heavy tail buffet loading. Also, a pitching moment that is associated with the leading separation has been found to be undesirable from the standpoint of controlling the airplane. Additionally, since this device is a critical flight control component it must have a redundant mechanical system to ensure operability in the event of failure of the main system, thus resulting in an increase in mechanical complexity and weight. Further drawbacks include the necessity for its attachment to primary aircraft structure, anti-icing considerations, and increased drag.
Another prior art device following the second approach is a body vane that is positioned on the side of the fuselage close to the wing leading edge. Since a vane in this position is within an area of very high airflow speed relative to free stream velocity, the vane has to be aligned with the airflow during cruise in order to minimize its drag effect and also must be attached to primary aircraft structure due to the forces imposed upon it. Since the vane may require a different setting for climbout, it would require a drive mechanism to control the setting of the vane. Furthermore, establishing the correct aerodynamic position from wind tunnel testing is unlikely, and easy variation of position in flight may be difficult to attain because of the required support structure.